The present invention relates to a gas turbine engine, and more particularly to airflow from a diffuser section of a tip turbine engine to cool various sections thereof.
An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan, a compressor, a combustor, and an aft turbine all located along a common longitudinal axis. A compressor and a turbine of the engine are interconnected by a shaft. The compressor is rotatably driven to compress air entering the combustor to a relatively high pressure. This pressurized air is then mixed with fuel in a combustor and ignited to form a high energy gas stream. The gas stream flows axially aft to rotatably drive the turbine which rotatably drives the compressor through the shaft. The gas stream is also responsible for rotating the bypass fan. In some instances, there are multiple shafts or spools. In such instances, there is a separate turbine connected to a separate corresponding compressor through each shaft. In most instances, the lowest pressure turbine will drive the bypass fan.
Although highly efficient, conventional turbofan engines operate in an axial flow relationship. The axial flow relationship results in a relatively complicated elongated engine structure of considerable longitudinal length relative to the engine diameter. This elongated shape may complicate or prevent packaging of the engine into particular applications.
A recent development in gas turbine engines is the tip turbine engine. Tip turbine engines locate an axial compressor forward of a bypass fan which includes hollow fan blades that receive airflow from the axial compressor therethrough such that the hollow fan blades operate as a centrifugal compressor. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor and ignited to form a high energy gas stream which drives the turbine integrated onto the tips of the hollow bypass fan blades for rotation therewith as generally disclosed in U.S. Patent Application Publication Nos.: 20030192303; 20030192304; and 20040025490.
The tip turbine engine provides a thrust to weight ratio equivalent to conventional turbofan engines of the same class within a package of significantly shorter length.
The tip turbine engine utilizes a fan-turbine rotor assembly which integrates a turbine onto the outer periphery of a diffuser section of the bypass fan. The diffuser turns and diffuses the airflow from the radial core airflow passage toward an annular combustor. Airflow separation and turbulence may occur as the airflow is turned from the radial core airflow passage toward an axial airflow direction which may reduce airflow to the annular combustor and reduce engine efficiency.
Accordingly, it is desirable to provide a diffuser for a fan-turbine rotor assembly of a tip turbine engine which minimizes airflow separation when passing therethrough.